This invention relates generally to gas turbine engines, and in particular, to a cooled flow path surface region on a turbine airfoil trailing edge.
This application references co-pending applications assigned to the assignee of the present invention, which are identified as Ser. No. 09/707,024 entitled xe2x80x9cTranspiration Cooling in Thermal Barrier Coating,xe2x80x9d; U.S. patent application Ser. No. 09/707,027 filed Nov. 6, 2000, now U.S. Pat. No. 6,375,425 issued Apr. 23, 2002; entitled xe2x80x9cMulti-layer Thermal Barrier Coating with Integrated Cooling System,xe2x80x9d; now U.S. application Ser. No. 09/707,024 filed Nov. 6, 2000; entitled xe2x80x9cDirectly Cooled Thermal Barrier Coating Systemxe2x80x9d now U.S. application Ser. No. 09/707,023 filed Nov. 6, 2002, entitled xe2x80x9cMethod For Creating Structured Porosity In Thermal Barrier Coating,xe2x80x9d now U.S. application Ser. No. 09/777,430 filed Feb. 6, 2001, the contents of which are incorporated herein by reference.
In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary-type compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on a shaft. The flow of gas turns the turbine, which turns the shaft and drives the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of combustion gases may exceed 3,000xc2x0 F., considerably higher than the melting temperatures of the metal parts of the engine which are in contact with these gases. Operation of these engines at gas temperatures that are above the metal part melting temperatures is a well established art, and depends in part on supplying a cooling fluid to the outer surfaces of the metal parts through various methods. Metal parts of these engines that are particularly subject to high temperatures, and thus require particular attention with respect to cooling, are, for example, the metal parts located aft of the combustor including high pressure turbine airfoils, such as exhaust nozzles and blades.
The hotter the turbine inlet gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine inlet gas temperature. However, the maximum temperature of the turbine inlet gases is normally limited by the materials used to fabricate the components downstream of the combustors such as the vanes and the blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to 2100xc2x0-2200xc2x0 F.
The metal temperatures can be maintained below melting levels with current cooling techniques by using a combination of improved active cooling designs and thermal barrier coatings (TBCs). For example, with regard to the metal blades and vanes employed in aircraft engines, some cooling is achieved through convection by providing passages for flow of cooling air from the compressor internally within the blades so that heat may be removed from the metal structure of the blade by the cooling air. Such blades have intricate serpentine passageways within the structural metal forming the cooling circuits of the blade.
Small internal orifices have also been devised to direct this circulating cooling air directly against certain inner surfaces of the airfoil to obtain cooling of the inner surface by impingement of the cooling air against the surface, a process known as impingement cooling. In addition, an array of small holes extending from a hollow core through the blade shell can provide for bleeding cooling air through the blade shell to the outer surface where a film of such air can protect the blade from direct contact with the hot gases passing through the engines, a process known as film cooling.
In another approach, a thermal barrier coating (TBC) is applied to the turbine blade component, which forms an interface between the metallic component and the hot gases of combustion. The TBC includes a ceramic coating that is applied to the external surface of metal parts to impede the transfer of heat from hot combustion gases to the metal parts, thus insulating the component from the hot combustion gas. This permits the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component.
TBCs include well-known ceramic materials, for example, yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do not adhere well directly to the superalloys used as substrate materials. Therefore, an additional metallic layer, a bond coat, is placed between the substrate and the TBC. The bond coat may be made of a nickel-containing overlay alloy, such as a MCrAlX, or other composition more resistant to environmental damage than the substrate, or alternatively, the bond coat may be a diffusion nickel aluminide or platinum aluminide whose surface oxidizes to form a protective aluminum oxide scale that provides improved adherence to the ceramic top coatings. The bond coat and overlying ceramic TBC are frequently referred to as a thermal barrier coating system.
Improved environmental resistance to destructive oxidation and hot corrosion is desirable. In addition, the alloying elements of the bond coat interdiffuse with the substrate alloy, changing the composition of the protective outer layer so that the walls of the turbine airfoils are consumed. This loss of material reduces the load carrying capability of the airfoil, thereby limiting blade life. This interdiffusion can also reduce environmental resistance of the coating. This interdiffusion and its adverse effects can be reduced by controlling the temperature of the component in the region of the bond coat/substrate interface.
In previous designs, the bond coat temperature limit has been critical to the TBC""s life and has had an upper limit of about 2100xc2x0 F. Once the bond coat exceeds this temperature, the thermal barrier coating system will quickly deteriorate, due to high temperature mechanical deformation and accelerated oxidation as well as a more rapid interdiffusion of elements between the bond coat and the underlying substrate alloy. The thermal barrier coating system can separate from the substrate exposing the underlying superalloy component to damage from the hot gasses.
Even with the use of advanced cooling designs and thermal barrier coatings, it is also desirable to decrease the requirement for cooling air, because reducing the demand for cooling air also contributes to improving overall engine operating efficiency. One way to achieve such a reduction is to improve the cooling of the metal parts immediately adjacent to their outer surfaces, which typically are exposed to the highest gas temperatures.
The trailing edge of high-pressure turbine airfoils, including nozzles and blades, typically require active cooling. Two types of trailing edge cooling are commonly used in current practice. The first type uses centerline convection cooling holes. This design requires a thicker trailing edge and, therefore, has more trailing edge blockage and lower aerodynamic efficiency, but has better cooling efficiency. The second type uses pressure side bleed film cooling slots/holes. This design permits the use of a thinner trailing edge and, therefore, has less trailing edge blockage and higher aerodynamic efficiency, but has lower cooling efficiency due to quick dissipation of the cooling film.
Thus, there is a need for a cooling design that can accept a thinner trailing edge of a turbine airfoil for better aerodynamic efficiency, yet still utilize the more effective convection cooling, rather than film cooling. In this manner, the environmental resistance and long-term stability of the thermal barrier coating system is improved and higher engine efficiencies can be obtained. The present invention fulfills this need, and further provides related advantages.
The present invention provides active convection cooling through micro channels within or adjacent to a bond coat layer applied to the trailing edge of a turbine engine high pressure airfoil. When placed adjacent to or within a porous TBC, the micro channels additionally provide transpiration cooling through the porous TBC. The micro channels communicate directly with at least one cooling circuit contained within the blade or vane from which they receive cooling air, thereby providing direct and efficient cooling for the bond coat layer. Because the substrate includes an actively cooled flow path surface region that can reduce the cooling requirement for the substrate, the engine can run at a higher firing temperature without the need for additional cooling air, achieving a better, more efficient engine performance.
In one embodiment, the present invention provides for an actively cooled flow path surface region of a gas turbine airfoil trailing edge comprising at least one trailing edge slot located within the airfoil substrate material, the slot having a first and second end. The first end terminates in a plenum located within the trailing edge, the second end communicates with a cooling fluid supply. At least one micro channel communicates with the plenum. The micro channel is located in a generally axial direction of the engine and parallel to the tip portion of the airfoil, along a surface of a first material applied over the substrate, such as a bond coat, and is substantially parallel to the surface of the substrate. A ceramic thermal barrier coating (TBC) overlies the first material and the micro channel.
In other preferred embodiments, the location of the micro channel may be, for example, at the substrate/bond coat interface, or it may be entirely within the TBC layer.
The present invention also sets forth a process for actively cooling the flow path surface region of a gas turbine engine airfoil trailing edge comprising the steps of casting an airfoil with pressure side bleed film cooling slots; adding a metallic bond coat to the airfoil; forming grooves in the bond coat such that the grooves are structured, with at least one structured micro groove communicating with at least one cooling fluid supply contained within the airfoil; applying, using a shadowing technique, a TBC layer over the structured grooves, resulting in the formation of hollow micro channels for the transport of a cooling fluid, and, passing a cooling fluid through the micro channels.
In other embodiments, the structured grooves, and therefore the resulting micro channels are located, for example, within the airfoil substrate at the substrate/bond coat interface, or they are placed entirely within the TBC layer.
The present invention further comprises the cooled flow path surface region formed by the foregoing processes and the turbine airfoil with the patterned micro channels substantially parallel to the surface of the substrate for cooling the component.
An advantage of the present invention is the flow path surface region of the coated gas turbine component is actively cooled. By removing heat from this region, the integrity of the bond coat can be maintained at higher engine operating temperatures.
In one embodiment, the active convection cooling through the micro channels occurs within or adjacent to the bond coat layer, providing direct and efficient cooling for the bond coat layer. Since the substrate is covered with the bond coat layer, the cooling requirement for the substrate will also be reduced.
Another advantage of the present invention is that the actively cooled bond coat layer will allow engine components to run at higher operating temperatures to achieve a better engine performance.
Still another advantage is that cooling air diffusing through the TBC will further lower the TBC temperature, thereby improving the TBC""s thermal insulation efficiency on the pressure surface for the trailing edge bond coat and substrate.
The increased cooling efficiency provided by the cooling channel and coating arrangement allows for the design of thin trailing edges having high aerodynamic efficiency and durability at higher gas path temperatures.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying figures which illustrate, by way of example, the principles of the invention.